CN1107934A - Gas turbine vane - Google Patents
Gas turbine vane Download PDFInfo
- Publication number
- CN1107934A CN1107934A CN94117613.4A CN94117613A CN1107934A CN 1107934 A CN1107934 A CN 1107934A CN 94117613 A CN94117613 A CN 94117613A CN 1107934 A CN1107934 A CN 1107934A
- Authority
- CN
- China
- Prior art keywords
- leading edge
- edge portion
- gas turbine
- aerofoil
- intermediate portion
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
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Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
- B23P—METAL-WORKING NOT OTHERWISE PROVIDED FOR; COMBINED OPERATIONS; UNIVERSAL MACHINE TOOLS
- B23P6/00—Restoring or reconditioning objects
- B23P6/002—Repairing turbine components, e.g. moving or stationary blades, rotors
- B23P6/005—Repairing turbine components, e.g. moving or stationary blades, rotors using only replacement pieces of a particular form
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/284—Selection of ceramic materials
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Architecture (AREA)
- Ceramic Engineering (AREA)
- Materials Engineering (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A vane for the turbine section of a gas turbine has an airfoil portion with leading edge, center and trailing edge portions. The leading edge portion is attached to the center portion by a dove tail joint that allows the leading edge portion to slide in the radial direction with respect to the center portion while preventing movement in the axial and circumferential directions, thereby eliminating thermal stresses created by differential thermal expansion between the leading edge portion and the remainder of the vane. An opening in the vane inner shroud that is normally sealed by a closure plate allows the leading edge portion to be readily replaced in the event of damage. The leading edge portion may be formed from a ceramic material and need not be supplied with cooling air.
Description
The present invention relates to employed aerofoil in the turbo machine, say more truely, the present invention relates to use the aerofoil in the stationary guide blades of gas turbine turbine part.
Gas turbine uses a plurality of stationary guide blades, they are arranged on the circumference of gas turbine turbine part, this class turborotor be exposed to the height of emitting in the combustion parts pine for, be exposed to sometimes the corrosion and have in the gas of particulate, thereby bring many problems to guide vane, comprise that burn into is ablated and because temperature is too high ruptures with thermal stress.These problems have reduced the working life of guide vane, thereby increase the utilization ratio of operating cost, reduction steam turbine.
The leading edge portion that flows directly into the guide vane aerofoil owing to hot gas causes very high thermal conductivity, so burn into is ablated and breakage problem is usually serious at the leading edge place.Unfortunately, because guide vane generally adopts the integrated type cast structure, so the reparation of leading edge portion need be carried out difficulty and expensive welding.
A kind of way that in the past was intended to improve the guide vane life-span is that the leading edge portion at aerofoil forms cooling air channels, thereby reduces its temperature.Unfortunately, reduced the thermal performance of turbine from the use of this class cooling air of combustion parts shunting.The another kind of way of Shi Yonging was to make guide vane with the stupalith with good high temperature resistant, corrosion protection, anti-ablation property in the past., this class stupalith is easily crisp, and holds and can't stand the very high thermal stress that causes owing to guide vane each several part differential thermal expansion, particularly when startup and shutdown.
Therefore, catalogue of the present invention provide a kind of gas turbine guide vane, its leading edge portion is repaired easily, and the useful ceramics material makes, but can not cause very high stress because of differential thermal expansion.
In brief, this purpose of the present invention and other purpose are finished in a kind of gas turbine, and the turbine portion branch of this gas turbine comprises a rotating shaft, and row's rotor blade is housed in the rotating shaft, and row's stationary guide blades is arranged on the circumference of rotating shaft.Each guide vane is characterised in that an aerofoil that is made of leading edge portion, rear edge part and the intermediate portion between the front and rear edge part and leading edge portion is fixed to device on the intermediate portion slidably, wherein, the further feature of guide vane be contained in the shield on the aerofoil and be formed on leading edge portion and this shield between a radial clearance.
In conjunction with the accompanying drawings to only as an example the explanation of preferred embodiment, the present invention can be more readily understood by following.
Fig. 1 is the combustion parts of the gas turbine of employing stationary guide blades of the present invention and the longitudinal sectional drawing of turbine part;
Fig. 2 is the stereogram of guide vane shown in Figure 1;
Fig. 3 is the ground plan of the interior shield part of guide vane shown in Figure 2;
Fig. 4 is that IV among Fig. 3 is to sectional drawing;
Fig. 5 is the stereogram of the leading edge portion of guide vane shown in Figure 2;
Fig. 6 is the main body figure of Fig. 3 and leading edge portion cover plate shown in Figure 4;
Fig. 7 is that VII among Fig. 4 is to sectional drawing.
Referring to each accompanying drawing, Fig. 1 illustrates the combustion parts 1 of gas turbine.U.S. Patent No. 4,991 391(Kosinski) is expressed a kind of typical gas turbine, so this specification is reference with it all.Combustion parts 1 is made of inner casing 4 and shell 5, forms ring casing 7 between the inside and outside shell.Inner casing 4 surrounds the rotor 6 that is positioned at the center.Combustion parts 1 also comprises a plurality of conduits 3, and the high hot combustion gas that temperature is surpassed 1200 ℃ (2200) causes turbine part 2 from the firing chamber (not shown).
Turbine part 2 comprises the blade ring 12 that is supported on the shell 5.Blade ring 12 is supporting row's stationary guide blades 16, and they are arranged on the circumference of rotor 6.Be contained in row's rotor blade 10 on the wheel rim part 8 of rotor 6 and be positioned at the downstream of this row's guide vane 16.From the pressurized air of the compressor section (not shown) ring casing 7 of flowing through.Though the most of pressurized air from compressor section is used for providing air for the burning of firing chamber, a part of pressurized air forms cooling air 19.Cooling air 19 passes through blade ring 12 from the firing chamber shunting hole 14 flows into guide vanes 16.
As shown in Figure 2, each guide vane 16 is made of interior shield 20, outer shield 22 and the aerofoil between inside and outside shield 24, and also, two shields are contained in the two ends of aerofoil.Hangers 32 and 34 extends on the shield, and guide vane 16 is fixed on the blade ring 12, as shown in Figure 1.According to the present invention, each aerofoil 24 comprises that one is contained in the demountable upstream leading edge portion 26 on the aerofoil main body.This aerofoil main body comprises downstream rear edge part 28, and this part 28 and the intermediate portion 27 between the front and rear edge part are made of one.
As shown in Figure 5, the leading edge portion 26 of aerofoil is one to have streamlined anterior 45 elongated element.As shown in Figure 7, the shape on anterior 45 surface is made with the outer convex surface of suction surface that forms aerofoil 24 main bodys respectively and pressure side and the slyness of inner concave 54 and 56 and is linked.
As shown in Figure 5, on the back of leading edge portion 26, stretch out a projection 44.In a preferred embodiment, the cross section of this projection 44 is roughly T shape.T section is preferably by angled and each limit non-rectilinear constitutes, thereby forms the shape that the technician generally is referred to as " dovetail ", as shown in Figure 7.The projection of stretching out on intermediate portion 27 fronts forms the groove 46 of the respective shapes that matches with leading edge portion projection 44, thereby projection 44 and groove 46 are pinned mutually.
As shown in Figure 7, when projection 44 and groove 46 matched, the interlocking that they form engaged and makes leading edge portion 26 also promptly can move freely with respect to intermediate portion 27 on the direction vertical with rotor 6 center lines diametrically.But this projection/groove engages and makes leading edge portion 26 axially and on the circumferencial direction moving, thereby leading edge portion is fixedly attached on the remaining part of guide vane.
The structure that this leading edge portion 26 can be done to move radially makes leading edge portion to slide along intermediate portion 27, thereby causes the present invention and existing stator that two big advantages are relatively arranged.
The first, because leading edge portion 26 can be independent of the remaining part of aerofoil 24 and substantially freely do radial expansion and contraction, therefore eliminated otherwise the localized heat stress that can in aerofoil, cause because of the differential thermal expansion between leading edge portion 26 and the intermediate portion 27.Second, if combine with following a kind of device that will further specify by one of shield insertion leading edge portion 26, this structure that moves radially can be by radially skidding off old leading edge portion and radially slip into new leading edge portion and the leading edge portion of for example burn into sintering or fracture is more changed simply and easily, thereby needn't adopt the repair welding work of costliness.
As shown in Figure 4, have an opening 43 on the interior shield 20.The shape of this opening 43 is identical with leading edge portion 26, but big slightly, and opening 43 aligns with leading edge portion 26 diametrically, thereby can insert opening to leading edge portion and slide on the intermediate portion 27.Shown in Fig. 3,4 and 6, cover plate 36 is inserted into after installing leading edge portion 26 and seals this opening on the opening 43.As shown in Figure 4, cover plate 36 usefulness beads 46 are soldered on the internal surface of interior shield.If will change leading edge portion 26, need only grind off bead 46 and remove cover plate 36.
Can see the most clearly from Fig. 6, the plug 44 that shape of cross section and leading edge portion are complementary stretches out on cover plate 36, and it makes it correctly in place along intermediate portion 27 against the bottom surface of living leading edge portion.
Shown in Fig. 4 is further, on outer shield 22 internal surfaces dimple 50 is arranged.The same with opening 43, the shape of this dimple 50 is identical with leading edge portion 26, but it is big slightly, and align with leading edge portion diametrically, this structure makes the top of leading edge portion 26 and bottom be located in respectively in the inside and outside shield (22 and 22), thereby forms smooth continuous fairing ace at the two ends of the contiguous inside and outside shield of aerofoil.According to leading edge portion 26 employed materials and employed material of aerofoil remaining part and their relatively hot expansion coefficient, be preferably between the dimple 50 of the end face of leading edge portion and outer shield and leave radial clearance 52.This gap 52 can expand leading edge portion 26 with respect to intermediate portion 27, and don't causes interlock between inside and outside shield, thereby sponges differential thermal expansion.
According to the present invention, because leading edge portion 26 is not made one with the aerofoil main body, therefore the available material different with the aerofoil main body made.As discussed further below, this feature main advantages of having compared with the integrated type aerofoil, but particularly when the said structure with the leading edge portion radial expansion combines.
In the past, by from the shunting of the cooling air of combustion parts and the intermediate portion 27 and the rear edge part 28 of the aerofoil that makes it to flow through cool off centre and rear edge part.Therefore, as Fig. 4 and shown in Figure 7, the inside of intermediate portion 27 and rear edge part 28 is hollow because of having cooling air channels 38 and 39 respectively substantially.Opening 40 and 41 on the outer shield 22 makes cooling air 14 from ring casing shown in Figure 17 flow channels 38 and 39.The hole 30 and 31 of arranged radially shown in Figure 2 makes a part of cooling air flow out and flow on the pressure side 56 of aerofoil 24 from passage 38 and 39 respectively.In addition, passage 48 makes a part of cooling air 14 flow through the downstream of the trailing edge of aerofoil 24.
But different with intermediate portion and rear edge part according to the present invention, the leading edge portion 26 of aerofoil 24 can be made solid, does not also promptly have cooling air channels, as Fig. 4 and shown in Figure 7.Owing to do not use cooling air, therefore avoided because of using the decline of the thermal performance that the additional cooling air of this class causes in leading edge portion 26.Why possible this improvement is, is that leading edge portion 26 is to use such as silicon nitride Si because in a preferred embodiment
3The stupalith of N and so on is made.This class material has good heat-resisting quantity and corrosion resistance, thereby does not need cooling air.And, because the slip joint structure of above-mentioned projection/groove has been eliminated leading edge portion 26 and the intersubjective inhomogeneous radial expansion of aerofoil, therefore overcome the shortcoming that frangible stupalith can't absorb thermal stress, and this shortcoming is limiting the application on the guide vane in gas turbine of frangible stupalith so far always.In addition, leading edge portion 26 also useful ceramics add that the composite of metal carrier body makes.Intermediate portion 27 and rear edge part 28 useful ceramics materials or the cobalt or other metal alloy that are used for turborotor are traditionally made, for example the Ni-based albronze of salt tolerant acid.
Though the invention has been described with reference to the airfoil portion of turborotor in the gas turbine above, but the present invention also can be applicable in other aerofoil of turbo machine, for example the aerofoil of the rotor blade of the aerofoil of the fixed guide vane of the compressor section of gas turbine or steam turbine or gas turbine or steam turbine.Therefore, only otherwise deviate from the spirit or the essential characteristic of appended claim, the present invention can realize by other concrete form, and thereby should be as the criterion with the appended claim that indicates the scope of the invention, rather than more than state specification and be as the criterion.
Claims (9)
1, a kind of gas turbine comprises:
Turbine part (2) with rotatingshaft (6), be equipped with on this rotatingshaft row's rotor blade (100, and
Row's stationary guide blades (16) around described axle (6) along circumferential arrangement, each described guide vane (16) comprises airfoil portion (24), it is characterized in that, described aerofoil (24) comprise leading edge portion (36), rear edge part (28) and be positioned at described leading edge portion (26) and described rear edge part (28) between intermediate portion (27) and described leading edge portion (26) is fixed to device (44,46) on the described intermediate portion (27) slidably; Wherein, described guide vane (16) is further characterized in that a radial clearance that is contained in the shield (22) on the described aerofoil (24) and forms between described leading edge portion (26) and described shield (22).
2, by the described gas turbine of claim 1, it is characterized in that described slip fixing device (44,46) is further characterized in that this device can radially slide by described relatively intermediate portion (27) described leading edge portion (26).
3,, it is characterized in that described slip fixing device is further characterized in that the first interlocking part (44) and the second interlocking part (46) give prominence to respectively on described leading edge portion (26) and intermediate portion (27) by the described gas turbine of claim 1.
By the described gas turbine of claim 3, it is characterized in that 4, one of described interlocking part forms a circumferentially extending groove (46).
5, by the described gas turbine of claim 4, it is characterized in that described groove (46) has a cross section that is roughly T shape.
6, by the described gas turbine of claim 1, wherein, described slip fixing device is characterised in that one is formed on described leading edge portion (26) and engages (44,46) with swallow-tail form between the described intermediate portion (27).
7, by the described gas turbine of claim 1, it is characterized in that described leading edge portion (26) is made by stupalith.
8, by the described gas turbine of claim 7, it is characterized in that described intermediate portion (27) is made by metallic material.
9, by the described gas turbine of claim 7, it is characterized in that described leading edge portion (26) is made by the stupalith of making parent with metal.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US08/141,756 US5358379A (en) | 1993-10-27 | 1993-10-27 | Gas turbine vane |
| US141,756 | 1993-10-27 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN1107934A true CN1107934A (en) | 1995-09-06 |
Family
ID=22497089
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN94117613.4A Pending CN1107934A (en) | 1993-10-27 | 1994-10-27 | Gas turbine vane |
Country Status (5)
| Country | Link |
|---|---|
| US (1) | US5358379A (en) |
| JP (1) | JPH07180504A (en) |
| CN (1) | CN1107934A (en) |
| AU (1) | AU672922B2 (en) |
| CA (1) | CA2134420A1 (en) |
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| US3215511A (en) * | 1962-03-30 | 1965-11-02 | Union Carbide Corp | Gas turbine nozzle vane and like articles |
| GB1030829A (en) * | 1965-04-27 | 1966-05-25 | Rolls Royce | Aerofoil blade for use in a hot fluid stream |
| US3487805A (en) * | 1966-12-22 | 1970-01-06 | Satterthwaite James G | Peripheral journal propeller drive |
| US3430898A (en) * | 1967-05-01 | 1969-03-04 | Us Navy | Leading edge for hypersonic vehicle |
| US3650635A (en) * | 1970-03-09 | 1972-03-21 | Chromalloy American Corp | Turbine vanes |
| US3856434A (en) * | 1973-10-18 | 1974-12-24 | Westinghouse Electric Corp | Centrifugal fan wheel |
| US4305697A (en) * | 1980-03-19 | 1981-12-15 | General Electric Company | Method and replacement member for repairing a gas turbine engine vane assembly |
| US4326833A (en) * | 1980-03-19 | 1982-04-27 | General Electric Company | Method and replacement member for repairing a gas turbine engine blade member |
| JPS59215905A (en) * | 1983-05-25 | 1984-12-05 | Hitachi Ltd | Wear resistant bucket |
| US5044884A (en) * | 1989-09-05 | 1991-09-03 | Trustees Of The University Of Pennsylvania | Safety propeller |
-
1993
- 1993-10-27 US US08/141,756 patent/US5358379A/en not_active Expired - Lifetime
-
1994
- 1994-10-12 AU AU75755/94A patent/AU672922B2/en not_active Expired - Fee Related
- 1994-10-26 CA CA002134420A patent/CA2134420A1/en not_active Abandoned
- 1994-10-27 JP JP6289137A patent/JPH07180504A/en not_active Withdrawn
- 1994-10-27 CN CN94117613.4A patent/CN1107934A/en active Pending
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| CN109154194B (en) * | 2016-03-16 | 2022-03-01 | 诺沃皮尼奥内技术股份有限公司 | Repair component for a gas turbine bucket assembly and method for repairing a damaged bucket of a gas turbine bucket assembly |
| CN112901278A (en) * | 2021-01-29 | 2021-06-04 | 大连理工大学 | Turbine blade adopting buckle fixed ceramic armor |
| CN112901279A (en) * | 2021-01-29 | 2021-06-04 | 大连理工大学 | Turbine blade adopting bolt-fixed ceramic armor |
| CN112901279B (en) * | 2021-01-29 | 2022-03-29 | 大连理工大学 | Turbine blade adopting bolt-fixed ceramic armor |
| CN112901278B (en) * | 2021-01-29 | 2022-03-29 | 大连理工大学 | Turbine blade adopting buckle fixed ceramic armor |
| CN114109519A (en) * | 2022-01-28 | 2022-03-01 | 中国航发沈阳发动机研究所 | Guide blade based on metal base band ceramic mosaic block |
Also Published As
| Publication number | Publication date |
|---|---|
| CA2134420A1 (en) | 1995-04-28 |
| JPH07180504A (en) | 1995-07-18 |
| US5358379A (en) | 1994-10-25 |
| AU7575594A (en) | 1995-05-18 |
| AU672922B2 (en) | 1996-10-17 |
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